Gas turbine engine inner barrel

ABSTRACT

A gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core. The gas turbine engine further comprises a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is based upon and claims the benefit of priority fromGB Patent Application No. GB 1806564.9, filed on 23 Apr. 2018, theentire contents of which are herein incorporated by reference.

BACKGROUND Technical Field

The present disclosure relates to an inner barrel for a gas turbineengine, the inner barrel bridging an inner ring from which outer guidevanes extend across a bypass duct of the engine and an inner cowl whichprovides an aerodynamic fairing surrounding the engine core.

Description of the Related Art

In a turbofan gas turbine engine, a propulsive fan generates twoairflows, one which passes through the core engine and one which passesthrough a surrounding bypass duct. Behind the fan in the bypass duct isa circumferential row of outer guide vanes which straighten out thebypass airflow from the fan. These vanes extend radially outwards froman inner ring which is a rigid structure defining a radially innersurface of the bypass duct.

Rearwardly of the plane of the outer guide vanes, the core engine issurrounded by an aerodynamic fairing called an inner cowl. This fairingalso defines a radially inner surface of the bypass duct, and typicallycomprises door sections that can be opened to allow maintenance accessto the core engine.

An interface structure if therefore needed to bridge the inner ring andthe inner cowl. As the inner ring is typically the responsibility of theengine manufacturer, while the configuration of the inner cowl can bethe responsibility of the airframer, this structure can take onadditional importance.

SUMMARY

According to a first aspect there is provided a gas turbine engine foran aircraft comprising:

an engine core comprising a compressor, a combustor, a turbine, and acore shaft connecting the turbine to the compressor;

a fan located upstream of the engine core, the fan comprising aplurality of fan blades, the fan generating a core airflow which entersthe engine core and a bypass airflow which flows through a bypass ductsurrounding the engine core;

a circumferential row of outer guide vanes located in the bypass ductrearwards of the fan, the outer guide vanes extending radially outwardlyfrom an inner ring which defines a radially inner surface of the bypassduct;

an inner cowl which provides an aerodynamic fairing surrounding theengine core, the inner cowl being rearwards of and axially spaced fromthe inner ring, and including one or more door sections which areopenable to enable maintenance access to the engine core; and

an inner barrel which surrounds the engine core and bridges the innerring and the inner cowl, the inner barrel having a circumferentiallyextending rear edge which provides an engagement formation forengagement with the door sections when they are closed.

The inner barrel thus not only bridges the inner ring and the innercowl, but also provides the engagement formation so that when the doorsections are fully closed a smooth aeroline can be formed along theradially inner surface of the bypass duct.

Optional features of the present disclosure will now be set out. Theseare applicable singly or in any combination with any aspect of thepresent disclosure.

The inner barrel may be formed as a unitary structure with the innerring. This can reduce the total weight of the engine and avoid a needfor fasteners joining the inner barrel to the inner ring.

The inner barrel may have openings for transmission of servicestherethrough. For example, the services can include air diverted fromthe bypass airflow for use in turbine case cooling and/or compressed airbled from the compressor for use in aircraft.

The gas turbine engine may further comprise an electronic unit such asan electronic core data controller which controls the engine core and/orperforms health monitoring of the engine core. Conveniently, theelectronic unit can be mounted to the inner barrel.

The gas turbine engine may further comprise a surface cooler for coolingengine fluid using the bypass air flow. Conveniently, the surface coolercan be mounted to the inner barrel. For example, the surface cooler canbe for cooling an integrated drive generator used to extract electricalpower from the engine core.

The inner barrel may have one or more drainage apertures for allowingdrainage of liquid from inside the engine barrel.

The inner barrel may have one or more ventilation holes for ventilatingthe engine core.

Conveniently, the engagement formation for engagement with the doorsections when they are closed may be a circumferentially extendingV-groove. The door sections can then have correspondingcircumferentially extending features which engage with the V-groove whenthe doors are closed.

The rear edge of the inner barrel may be forward of the combustor.

An outer surface of the inner barrel typically defines a radially innersurface of the bypass duct. For example, this outer surface can be asurface of a main structural element (e.g. engine fire zone boundary) ofthe barrel, and/or it can be one or more in-fill panels located inoutwardly facing recesses formed in the barrel, and/or it can be one ormore fairings supported by a main structural element of the barrel.

It is possible for the inner barrel to be formed as a unitary piece.However, another option is for the inner barrel to be formed from pluralseparate barrel portions. In particular, the inner barrel may be formedas two half barrels located on respective opposite sides of the engine.The half barrels can be spaced apart at the top of the engine by amounting pylon for mounting the engine to an airframe. For example, themounting pylon can join to the engine core at a fixture located at topdead centre behind the inner ring.

Additionally or alternatively, when the inner barrel is formed as twohalf barrels on respective opposite sides of the engine, the inner cowlmay have two door sections located on the respective opposite sides ofthe engine, each door section being pivotably openable about arespective pivot line which extends from front to back e.g. along arespective side of the above-mentioned mounting pylon. The two halfbarrels may then be spaced apart at the bottom of the engine by a keelbeam which extends rearwardly from the inner ring at bottom dead centrethereof to provide latching formations rearward of the inner barrel forlatching to lower edges of the door sections when they are closed.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹K⁻¹/(ms⁻¹)²), The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹ s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹ s or80 Nkg⁻¹ s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55 deg C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 shows schematically a perspective view from the rear of an enginewithout its nacelle and without its inner cowl;

FIG. 5 shows a side view of a half barrel of an inner barrel of theengine of FIG. 4;

FIG. 6 shows a cross-section through the inner barrel of FIG. 5;

FIG. 7 shows a cross-section through a variant inner barrel; and

FIG. 8 shows at top the inner barrel of FIGS. 4 to 6 attached byfasteners to an inner ring of the engine, and at bottom a variant of theinner barrel formed as a unitary structure with the inner ring.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine is mounted to an airframe, e.g. under a wing, by amounting pylon 46. The engine 10 comprises an air intake 12 and apropulsive fan 23 that generates two airflows: a core airflow A and abypass airflow B. The gas turbine engine 10 comprises a core 11 thatreceives the core airflow A. The engine core 11 comprises, in axial flowseries, a low pressure compressor 14, a high-pressure compressor 15,combustion equipment 16, a high-pressure turbine 17, a low pressureturbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gasturbine engine 10 and defines a bypass duct 22 and a bypass exhaustnozzle 18. The pylon 46 forms an upper bifurcation in the bypass ductwhere it traverses the duct to join to the engine core 11. The bypassairflow B flows through the bypass duct 22, where it is straightened bya row of outer guide vanes 40 before exiting the bypass exhaust nozzle18. Rearward of the outer guide vanes 40, the engine core 10 issurrounded by an inner cowl 41 which provides an aerodynamic fairingdefining an inner surface of the bypass duct 22. The fan 23 is attachedto and driven by the low pressure turbine 19 via a shaft 26 and anepicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22meaning that the flow through the bypass duct 22 has its own nozzle thatis separate to and radially outside the core engine nozzle 20. However,this is not limiting, and any aspect of the present disclosure may alsoapply to engines in which the flow through the bypass duct 22 and theflow through the core 11 are mixed, or combined, before (or upstream of)a single nozzle, which may be referred to as a mixed flow nozzle. One orboth nozzles (whether mixed or split flow) may have a fixed or variablearea. In some arrangements, the gas turbine engine 10 may not comprise agearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 shows schematically a perspective view from the rear of theengine 10 with its nacelle 21, inner cowl 41 and pylon 46 removed. A fancase 42 defines an outer surface of the bypass duct 22 and towards therear of the fan case an inner ring 44 defines an inner surface of thebypass duct 22. The outer guide vanes 40 extend radially from the innerring to the fan case, and the engine core 11 projects rearwardly fromthe plane of the outer guide vanes. A fixture 45 located at top deadcentre behind the inner ring provides a connection point for themounting pylon 46 which mounts the engine to the airframe.

The inner cowl 41 can be formed as two door sections, one on either sideof the engine 10, with each door section being pivotably openable abouta respective pivot line 54 which extends from front to back along thatdoor section's side of the pylon 46. This allows the door sections to beswung upwards and away from the engine core 11 for maintenance access.Conveniently, the top edges of the door sections can form the pivotlines. A keel beam 48 can also be provided, the keel beam extendingrearwardly from bottom dead centre of the inner ring 44 to providelatching formations for latching to lower edges of the door sectionswhen they are closed.

The inner cowl 41 is axially spaced from the inner ring 44. An innerbarrel 47 bridges this space and provides a gas-washed fireproofstructural panel with amalgamated fire sealing, drainage, and mechanicalsupport interface features. The inner barrel 47 is in two halves 47 a,47 b located on opposite sides of the engine. Each half barrel extendscircumferentially on its side of the engine from the fixture 45 to thekeel beam 48. In other variants, however, the inner barrel can be formedfrom more than two components, or can be formed as a single, continuouscomponent (although in that case, the fixture 45 has to be moved oradapted to allow the barrel to continue through top dead centre, and/orthe keel beam 48 has to be adapted or removed to allow the barrel tocontinue through bottom dead centre). Typically, the inner ring extendsonly a limited length in the axial direction, e.g. such that its rearedge is forward of the combustor of the engine core 11. This isconsistent with providing an overall axially compact engine.

The inner barrel can provide some or all of the followingfunctionalities:

-   -   (1) A zone compliant (fire and fluid) boundary between core and        bypass zones of the engine.    -   (2) An interface to the inner cowl 41, such as circumferentially        extending V-groove 49 for engagement, and hence position and        load sharing, with the door sections of the inner cowl 41 when        they are closed.    -   (3) An inner aeroline 51 of the bypass duct 22 along its axial        station.    -   (4) Ventilation features for core zone ventilation.    -   (5) Drainage surfaces and apertures for core zone drainage.    -   (6) A fire zone compliant interface for turbine case cooling        (TCC) air ingestion from the bypass duct 22.    -   (7) A fire zone compliant interface(s) for compressed air bleed        exit from one of the compressors and optionally for other bleed        air systems.    -   (8) Structural support and positioning features for mounting an        electronic unit such as an electronic core data controller which        controls the engine core 11 and/or performs health monitoring of        the engine core.    -   (9) Structural support and positioning features for mounting a        surface cooler for cooling engine fluid using the bypass air        flow. For example, the cooled engine fluid can be used to cool        an integrated drive generator for extracting electrical power        from the engine core 11.    -   (10) Acoustic treatment for noise attenuation.

FIG. 5 shows a side view of the half barrel 47 a and illustratesfeatures of the inner barrel providing some of these functionalities,such as the V-groove 49. FIG. 6 shows a cross-section through the innerbarrel, with the core and bypass zones of the engine indicated on eitherside of the zone compliant boundary 50 provided by the barrel. The inneraero line 51 of the bypass duct 22 can then be provided by a fairingsupported by the zone compliant boundary. However, FIG. 7 shows across-section through a variant inner barrel in which the zone compliantboundary 50 also forms the inner aeroline of the bypass duct. In thiscase, any bypass-side components mounted to the barrel can be located inlocal recesses 56 formed on that side of the barrel, as shown in theinset to FIG. 7.

The forward edge of the inner barrel can have fastening features, suchas a circumferential row of fastener (e.g. screw and/or bolt) holes 52for fastening the barrel to the inner ring 44, as shown in the topcross-section of FIG. 8. However, another option, shown in the bottomcross-section of FIG. 8, is to form the inner barrel as a unitarystructure with the inner ring, thereby reducing the total weight of theengine and avoiding a need for fasteners.

The zone compliant boundary 50 of the inner barrel can be fabricatedfrom a metal sheet, which can then be forged to a ring providing theV-groove 49 and another ring providing a flange for the bolt holes 52.However, another option is to form at least the zone compliant boundarythe barrel from composite material, e.g. by a laying up process.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

The invention claimed is:
 1. A gas turbine engine for an aircraftcomprising: an engine core comprising a compressor, a combustor, aturbine, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades, the fan generating a core airflow which enters the enginecore and a bypass airflow which flows through a bypass duct surroundingthe engine core; a circumferential row of outer guide vanes located inthe bypass duct rearwards of the fan, the outer guide vanes extendingradially outwardly from an inner ring which defines a portion of aradially inner surface of the bypass duct; an inner cowl which providesan aerodynamic fairing surrounding the engine core, the inner cowl beingrearwards of and axially spaced from the inner ring; and an inner barrelwhich surrounds the engine core and bridges the inner ring and the innercowl, the inner barrel having a circumferentially extending rear edgewhich provides a circumferentially extending v-groove, wherein the innerbarrel is formed as two half barrels spaced apart and located onrespective opposite sides of the gas turbine engine, wherein the innerbarrel comprises a local recess formed on a side of the inner barrelfacing the bypass duct, wherein the local recess comprises a pair ofsidewalls formed by a radially outer surface of the inner barrel, thesidewalls extending radially inward away from the bypass duct.
 2. A gasturbine engine according to claim 1, wherein the inner barrel is formedas a unitary structure with the inner ring.
 3. A gas turbine engineaccording to claim 1, wherein the rear edge of the inner barrel isforward of the combustor.
 4. A gas turbine engine according to claim 1,wherein an outer surface of the inner barrel defines another portion ofthe radially inner surface of the bypass duct.
 5. A gas turbine engineaccording to claim 1, wherein the half barrels are spaced apart at a topof the gas turbine engine by a mounting pylon for mounting the gasturbine engine to an airframe.
 6. A gas turbine engine according toclaim 5, wherein: the two half barrels are spaced apart at a bottom ofthe gas turbine engine by a keel beam which extends rearwardly from theinner ring at bottom dead centre thereof.
 7. A gas turbine engineaccording to claim 1, further comprising a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft.
 8. The gasturbine engine according to claim 1, wherein: the turbine is a firstturbine, the compressor is a first compressor, and the core shaft is afirst core shaft; the engine core further comprises a second turbine, asecond compressor, and a second core shaft connecting the second turbineto the second compressor; and the second turbine, second compressor, andsecond core shaft are arranged to rotate at a higher rotational speedthan the first core shaft.
 9. The gas turbine engine according to claim1, wherein an axially extending surface of the inner barrel is disposedradially inward of an immediately adjacent portion of the inner ringupstream of the inner barrel.
 10. A gas turbine engine for an aircraftcomprising: an engine core comprising a compressor, a combustor, aturbine, and a core shaft connecting the turbine to the compressor; afan located upstream of the engine core, the fan comprising a pluralityof fan blades, the fan generating a core airflow which enters the enginecore and a bypass airflow which flows through a bypass duct surroundingthe engine core; a circumferential row of outer guide vanes located inthe bypass duct rearwards of the fan, the outer guide vanes extendingradially outwardly from an inner ring which defines a portion of aradially inner surface of the bypass duct; an inner cowl which providesan aerodynamic fairing surrounding the engine core, the inner cowl beingrearwards of and axially spaced from the inner ring; and an inner barrelwhich surrounds the engine core and bridges the inner ring and the innercowl, the inner barrel having a circumferentially extending rear edgewhich provides a circumferentially extending v-groove, wherein the innerbarrel is formed as two half barrels spaced apart and located onrespective opposite sides of the gas turbine engine, wherein an axiallyextending surface of the inner barrel is disposed radially inward of animmediately adjacent portion of the inner ring upstream of the innerbarrel, wherein the axially extending surface of the inner barrel is aradially outer surface of the inner barrel facing the bypass duct.
 11. Agas turbine engine for an aircraft comprising: an engine core comprisinga compressor, a combustor, a turbine, and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of fan blades, the fan generating a coreairflow which enters the engine core and a bypass airflow which flowsthrough a bypass duct surrounding the engine core; a circumferential rowof outer guide vanes located in the bypass duct rearwards of the fan,the outer guide vanes extending radially outwardly from an inner ringwhich defines a portion of a radially inner surface of the bypass duct;an inner cowl which provides an aerodynamic fairing surrounding theengine core, the inner cowl being rearwards of and axially spaced fromthe inner ring; and an inner barrel which surrounds the engine core andbridges the inner ring and the inner cowl, the inner barrel having acircumferentially extending rear edge which provides a circumferentiallyextending v-groove, wherein the inner barrel is formed as two halfbarrels spaced apart and located on respective opposite sides of the gasturbine engine, wherein the inner barrel comprises a local recess formedon a side of the inner barrel facing the bypass duct, wherein the localrecess does not extend across an entire circumference of the innerbarrel.